Dynamic explicit guidance method, device, medium and product based on Newton iteration

Abstract

The present invention discloses a dynamic explicit guidance method, device, medium, and product based on Newton iteration, relating to the field of guidance technology. Newton iteration is used to update the main quantity of thrust direction, the change in thrust direction, and the engine shutdown time; numerical integration and Newton iteration are used to update the position to be increased, the speed to be increased, the first position control direction, and the second position control direction; whether the iteration has converged is determined to obtain a first determination result; if converged, the current moment is updated to the next moment, and the function "obtaining the flight status of the rocket" is returned; otherwise, the updated engine shutdown time and the updated thrust direction are used to control the operation of the rocket, the current moment is updated to the next moment, the function "obtaining the flight status of the rocket" is returned, and it is determined whether the rocket has reached the target orbit; if so, the guidance is terminated; if not, the current moment is updated to the next moment, and the function "obtaining the flight status of the rocket" is returned. The present invention improves the applicability of dynamic explicit guidance.

Description

Technical Field

[0001] The present invention relates to the field of guidance technology, and in particular to a dynamic explicit guidance method, device, medium and product based on Newton iteration. Background Art

[0002] A key development direction for intelligent control in aerospace is enhancing the adaptability of launch vehicle guidance methods to flight missions and non-fatal failures. Inserting into a CMO is one of the most challenging tasks for ascent-stage guidance. Compared to low-Earth orbit (LEO), the CMO has a higher energy, requiring a longer propulsion time to meet target velocity constraints. The resulting long propulsion arc complicates accurate thrust prediction and guidance planning. Furthermore, the CMO has a large eccentricity. Unlike circular or near-circular orbits, the velocity amplitude at the insertion point varies with its position, creating a significant coupling effect and posing challenges for guidance methods to accurately correct the insertion point. Therefore, directly applying guidance methods designed for LEO to CMO insertion missions can lead to reduced accuracy and fuel optimality, and even iterative divergence.

[0003] Powered explicit guidance is a widely used guidance method for rocket ascent phases. It originates from the linear tangent guidance method developed by NASA for the Space Shuttle. The Space Shuttle, during some of its exoatmospheric missions, has a very low thrust-to-weight ratio and a very long flight time. This does not satisfy the short propulsion arc assumption of previous iterative guidance methods. Consequently, iterative guidance applications on the Space Shuttle experienced performance degradation and even iterative divergence. To address this, NASA developed various guidance methods and selected the linear tangent guidance method as the foundation for the development of the powered explicit guidance method. Compared to previous iterative guidance methods, powered explicit guidance boasts greater mission adaptability by employing linear thrust direction instead of linear program angle assumptions, flight time instead of flight time, and an analytical gravity field instead of a constant gravity field. This improvement allows it to handle diverse exoatmospheric missions. NASA has subsequently implemented numerous improvements to powered explicit guidance. NASA's latest Space Launch System (SLS) continues to utilize the powered explicit guidance method developed for the Space Shuttle during ascent phases. However, during the upper stage's orbital insertion mission, the long propulsion arc causes long-term nonlinear changes in the thrust direction, and the small-angle Taylor expansion method used in the powered explicit guidance cannot accurately predict the thrust effect; in addition, the large eccentricity of the Earth-Moon transfer orbit makes the position and velocity of the insertion point highly coupled, which may cause the iterative loop of the velocity to be increased in the powered explicit guidance to become unstable. As a result, the powered explicit guidance has experienced performance degradation and even iterative divergence, and cannot work normally in some Earth-Moon transfer mission scenarios. Summary of the Invention

[0004] The purpose of the present invention is to provide a dynamic explicit guidance method, device, medium and product based on Newton iteration, thereby improving the applicability of dynamic explicit guidance.

[0005] To achieve the above object, the present invention provides the following solutions:

[0006] A dynamic explicit guidance method based on Newton iteration, comprising:

[0007] Obtaining initial values ​​of orbital elements and iterative variables of the target orbit of the rocket; the orbital elements include: semi-major axis, eccentricity, orbital inclination, right ascension of ascending node, and argument of perigee; and the iterative variables include: position to be increased, velocity to be increased, first position control direction, second position control direction, main quantity of thrust direction, change in thrust direction, and engine shutdown time;

[0008] Obtaining the flight status of the rocket; the flight status includes: position, speed, mass, thrust magnitude and effective exhaust velocity;

[0009] At the current moment, using Newton iteration, based on the flight status, the main quantity of the thrust direction, the change in the thrust direction, and the engine shutdown time are updated to obtain an updated main quantity of the thrust direction, an updated change in the thrust direction, and an updated engine shutdown time;

[0010] Using numerical integration and Newton iteration, based on the flight state, the updated main quantity of the thrust direction, the updated change in the thrust direction, the updated engine shutdown time, and the orbital elements, the position to be increased, the speed to be increased, the first position control direction, and the second position control direction are updated to obtain an updated position to be increased, an updated speed to be increased, an updated first position control direction, and an updated second position control direction;

[0011] Based on the updated engine shutdown time and convergence judgment inequality, judging whether the iteration has converged to obtain a first judgment result;

[0012] If the first judgment result is no, the current moment is updated to the next moment, and the process returns to "getting the rocket's flight status";

[0013] If the first judgment result is yes, outputting the updated engine shutdown time and the updated thrust direction to the control system of the rocket; the updated thrust direction is determined according to the updated main quantity of the thrust direction and the updated change quantity of the thrust direction;

[0014] The control system controls the operation of the rocket using the updated engine shutdown time and the updated thrust direction, updates the current moment to the next moment, returns to "obtaining the rocket's flight status", and determines whether the rocket has reached the target orbit based on the flight status and orbital elements, thereby obtaining a second determination result;

[0015] If the second judgment result is yes, then end the guidance;

[0016] If the second judgment result is no, the current moment is updated to the next moment, and the process returns to "getting the rocket's flight status".

[0017] Optionally, get the initial value of the rocket's iteration variable, including:

[0018] Using the least propellant as the performance indicator, trajectory optimization software is used to optimize the rocket trajectory before launch to determine the position curve, velocity curve, mass curve and thrust curve, position co-state curve, velocity co-state curve and engine shutdown time;

[0019] Based on the position curve, velocity curve, mass curve and thrust curve, position co-state curve, velocity co-state curve and engine shutdown time, the initial value of the rocket's iterative variable is obtained using the iterative variable calculation formula.

[0020] Optionally, the iterative variable calculation formula includes:

[0021] ;

[0022] ;

[0023] ;

[0024] ;

[0025] ;

[0026] ;

[0027] ;

[0028] in, For the position to be added; The engine shutdown time obtained by optimizing the rocket trajectory; For the time; For the The thrust of the moment; For the the quality of the moment; For the Velocity covariance at each moment; is the speed to be increased; Control direction for the first position; The position of the engine shutdown time obtained for optimizing the rocket trajectory; The speed of the engine shutdown time obtained for optimizing the rocket trajectory; Control direction for the second position; is the main quantity in the thrust direction; is the velocity co-state at the initial moment; is the change in thrust direction; The engine shutdown time is obtained by using the iterative variable calculation formula.

[0029] Optionally, using Newton iteration, updating the main quantity of the thrust direction, the change in the thrust direction, and the engine shutdown time based on the flight state to obtain an updated main quantity of the thrust direction, an updated change in the thrust direction, and an updated engine shutdown time, includes:

[0030] Constructing a first nonlinear equation based on the flight state; the first nonlinear equation is an equation about a main quantity of the thrust direction, a change in the thrust direction, and an engine shutdown time;

[0031] Solving the first nonlinear equation using Newton iteration to obtain an updated amount of a main quantity in the thrust direction, an updated amount of a change in the thrust direction, and an updated amount of an engine shutdown time;

[0032] updating the main quantity in the thrust direction based on the main quantity update amount in the thrust direction to obtain an updated main quantity in the thrust direction;

[0033] updating the change amount of the thrust direction based on the change amount update amount of the thrust direction to obtain an updated change amount of the thrust direction;

[0034] The engine shutdown time is updated based on the engine shutdown time update amount to obtain an updated engine shutdown time.

[0035] Optionally, using numerical integration and Newton iteration, based on the flight state, the updated main quantity of the thrust direction, the updated change in the thrust direction, the updated engine shutdown time, and the orbital elements, the position to be increased, the speed to be increased, the first position control direction, and the second position control direction are updated to obtain the updated position to be increased, the updated speed to be increased, the updated first position control direction, and the updated second position control direction, including:

[0036] Constructing a differential equation based on the flight state, the principal quantity of the updated thrust direction, and the change quantity of the updated thrust direction;

[0037] Applying the fourth-order Runge-Kutta method to solve the differential equation to obtain the predicted position and predicted velocity;

[0038] Constructing a second nonlinear equation based on the track element, the predicted position, and the predicted speed; wherein the second nonlinear equation is an equation about the corrected position and the corrected speed;

[0039] Using Newton iteration to solve the second nonlinear equation to obtain the corrected position and the corrected velocity;

[0040] Based on the corrected position and the corrected speed, the position to be increased, the speed to be increased, the first position control direction and the second position control direction are updated to obtain an updated position to be increased, an updated speed to be increased, an updated first position control direction and an updated second position control direction.

[0041] Optionally, the convergence judgment inequality is:

[0042] ;

[0043] in, The engine shutdown time is obtained by using the iterative variable calculation formula; The stored engine shutdown time.

[0044] Optionally, the thrust direction is calculated as:

[0045] ;

[0046] in, For the The thrust direction at the moment; is the main quantity in the thrust direction; For the time; is the change in thrust direction.

[0047] A computer device comprises: a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement any one of the above-mentioned dynamic explicit guidance methods based on Newton iteration.

[0048] A computer-readable storage medium stores a computer program, which, when executed by a processor, implements any of the above-mentioned dynamic explicit guidance methods based on Newton iteration.

[0049] A computer program product comprises a computer program, wherein when the computer program is executed by a processor, the computer program implements any one of the above-mentioned dynamic explicit guidance methods based on Newton iteration.

[0050] According to the specific embodiments provided by the present invention, the present invention discloses the following technical effects:

[0051] The present invention discloses a dynamic explicit guidance method, device, medium, and product based on Newton iteration. Aiming at the long propulsion arc and large eccentricity characteristics of Earth-Moon transfer conditions, a number of numerical methods, including Newton iteration and numerical integration, are introduced to improve the mission adaptability and iterative convergence of the guidance method. By combining the iterative process of existing dynamic explicit guidance methods with the high-precision characteristics of numerical methods, a rocket guidance method is obtained that can better adapt to conditions with long propulsion arcs and large eccentricity. It has advantages in attitude stability and fuel optimality and can adapt to long propulsion arc conditions with flight times of more than 1,500 seconds. It has engineering application value in missions such as manned lunar landings. BRIEF DESCRIPTION OF THE DRAWINGS

[0052] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0053] Figure 1 A schematic flow chart of a dynamic explicit guidance method based on Newton iteration provided in Example 1 of the present invention;

[0054] Figure 2 This is a diagram of the internal structure of a computer device. DETAILED DESCRIPTION

[0055] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0056] The purpose of the present invention is to provide a dynamic explicit guidance method, device, medium and product based on Newton iteration, aiming to improve the applicability of dynamic explicit guidance.

[0057] In order to make the above-mentioned objects, features and advantages of the present invention more obvious and easy to understand, the present invention is further described in detail below with reference to the accompanying drawings and specific embodiments.

[0058] Example 1

[0059] like Figure 1 As shown, the dynamic explicit guidance method based on Newton iteration in this embodiment includes:

[0060] Step 1: Obtain the orbital elements of the rocket's target orbit and the initial values ​​of the iteration variables.

[0061] Among them, the orbital elements include: semi-major axis, eccentricity, orbital inclination, right ascension of ascending node and argument of perigee; the iterative variables include: position to be increased, speed to be increased, control direction of the first position, control direction of the second position, main quantity of thrust direction, change in thrust direction and engine shutdown time.

[0062] Specifically, the semi-major axis of the target orbit , eccentricity , orbital inclination , right ascension of the ascending node and the argument of perigee These five orbital elements can uniquely determine the target orbit of a rocket. All five orbital elements are scalar quantities and are well-known concepts in the field of aerospace control.

[0063] Engine shutdown time Is a scalar, the position to be increased , speed to be increased , first position control direction , Second position control direction , the main quantity in the thrust direction and the change in thrust direction are all three-dimensional vectors.

[0064] The initial values ​​of the iterative variables are designed before the rocket is launched and loaded into the rocket's control system as launch parameters.

[0065] As an optional implementation, obtaining the initial value of the iteration variable of the rocket includes:

[0066] Step 101: Using the least propellant as a performance indicator, use trajectory optimization software to optimize the rocket trajectory before launch to determine the position curve, velocity curve, mass curve and thrust curve, position co-state curve, velocity co-state curve and engine shutdown time.

[0067] Specifically, general trajectory optimization software, such as GPOPS software commonly used in the field of aerospace control, is used for optimization design.

[0068] The position curve, velocity curve, mass curve, thrust curve, position co-state curve and velocity co-state curve are all functions of time. Position at the moment , No. The speed of time , No. Position covariance at a moment Hedi Velocity covariance at each moment are all three-dimensional vectors, Quality of the moment Hedi The push of time Both are scalar

[0069] Step 102: Based on the position curve, velocity curve, mass curve and thrust curve, position co-state curve, velocity co-state curve and engine shutdown time, the iterative variable calculation formula is used to obtain the initial value of the rocket's iterative variable.

[0070] As an optional implementation, the iterative variable calculation formula includes:

[0071] .

[0072] .

[0073] .

[0074] .

[0075] .

[0076] .

[0077] .

[0078] in, For the position to be added; The engine shutdown time obtained by optimizing the rocket trajectory; For the time; For the The thrust of the moment; For the the quality of the moment; For the Velocity covariance at each moment; is the speed to be increased; Control direction for the first position; The position of the engine shutdown time obtained for optimizing the rocket trajectory; The speed of the engine shutdown time obtained for optimizing the rocket trajectory; Control direction for the second position; is the main quantity in the thrust direction; is the velocity co-state at the initial moment; is the change in thrust direction; The engine shutdown time is obtained by using the iterative variable calculation formula.

[0079] Step 2: Get the rocket's flight status.

[0080] Among them, the flight status includes: position, speed, mass, thrust size and effective exhaust speed.

[0081] Specifically, location and speed are all three-dimensional vectors, with mass , thrust size and the engine's effective exhaust velocity Both are scalars. Position is the coordinate of the rocket's center of mass in the geocentric inertial coordinate system; speed For location Derivative with respect to time; mass Indicates the total mass of the rocket; thrust Indicates the total thrust of all engines of the rocket; effective exhaust velocity It represents the average effective exhaust velocity of all the rocket's engines.

[0082] Step 3: At the current moment, using Newton iteration, update the main quantity of thrust direction, the change in thrust direction, and the engine shutdown time based on the flight status to obtain the updated main quantity of thrust direction, the updated change in thrust direction, and the updated engine shutdown time.

[0083] As an optional implementation, step 3 includes:

[0084] Step 301: Construct a first nonlinear equation based on the flight state; the first nonlinear equation is an equation about the main quantity of the thrust direction, the change in the thrust direction, and the engine shutdown time.

[0085] Specifically, first, consider the second-order differential equation for rocket thrust acceleration:

[0086] .

[0087] in, is the second-order integral of the rocket thrust acceleration The first derivative of ; is the first-order integral of the rocket thrust acceleration The first derivative of ; and are all three-dimensional vectors.

[0088] Solving the integral form of the above differential equation, we get:

[0089] .

[0090] in, Based on 、 and Determine the second-order integral of the rocket thrust acceleration; Based on 、 and Determine the first-order integral of the rocket thrust acceleration.

[0091] Second, consider the following about 、 and The nonlinear equation, that is, the first nonlinear equation:

[0092] .

[0093] in, is a seven-dimensional vector, and its expression is:

[0094] .

[0095] in, is the transpose of the vector; For the the speed of the moment; For the Position at the moment The acceleration of gravity at .

[0096] Step 302: Newton iteration is used to solve the first nonlinear equation to obtain the main quantity update of the thrust direction, the change quantity update of the thrust direction, and the engine shutdown time update.

[0097] Specifically, Newton iteration is applied to solve the first nonlinear equation. The input of the first nonlinear equation is a seven-dimensional vector (3+3+1), and the output is also a seven-dimensional vector (1+1+3+1+1). Its gradient matrix (also called Jacobian matrix) is for:

[0098] .

[0099] Among them, the gradient matrix is a matrix with seven rows and seven columns, computed using numerical differences.

[0100] Solve the following about 、 and The linear equations of the thrust direction are obtained by , the change in thrust direction and engine shutdown time update :

[0101] .

[0102] Step 303: updating the main quantity in the thrust direction based on the main quantity update amount in the thrust direction to obtain an updated main quantity in the thrust direction.

[0103] Specifically, the update formula of the main quantity in the thrust direction is:

[0104] .

[0105] in, is the main quantity in the updated thrust direction; is the main quantity in the thrust direction before updating.

[0106] Step 304: updating the change amount of the thrust direction based on the updated change amount of the thrust direction to obtain an updated change amount of the thrust direction.

[0107] Specifically, the update formula for the change in thrust direction is:

[0108] .

[0109] in, is the change in thrust direction after update; is the change in thrust direction before updating.

[0110] Step 305: Update the engine shutdown time based on the engine shutdown time update amount to obtain an updated engine shutdown time.

[0111] Specifically, the update formula for the engine shutdown time is:

[0112] .

[0113] in, The updated engine shutdown time; The engine shutdown time before the update.

[0114] Step 4: Using numerical integration and Newton iteration, based on the flight state, the main quantity of the updated thrust direction, the change in the updated thrust direction, the updated engine shutdown time and the orbital elements, update the position to be increased, the speed to be increased, the first position control direction and the second position control direction to obtain the updated position to be increased, the updated speed to be increased, the updated first position control direction and the updated second position control direction.

[0115] As an optional implementation, step 4 includes:

[0116] Step 401: Construct a differential equation based on the flight state, the updated principal quantity of the thrust direction, and the updated change quantity of the thrust direction.

[0117] Specifically, consider the following differential equation:

[0118] .

[0119] in, For location The first derivative of ; For speed The first derivative of ; For quality The first derivative of ; is the first-order integral of the acceleration due to gravity The first derivative of ; is the second-order integral of the acceleration due to gravity The first derivative of ; is the gravitational acceleration, which is a three-dimensional vector with respect to the position The expression is:

[0120] .

[0121] in, is the Earth's gravitational constant, which is a scalar.

[0122] Step 402: Apply the fourth-order Runge-Kutta method to solve the differential equation to obtain the predicted position and predicted velocity.

[0123] Specifically, the fourth-order Runge-Kutta method is used to solve the differential equation, and the solution interval is set to , set the initial value of the differential equation to In the solution of the fourth-order Runge-Kutta method, the definition and Represents the predicted position and predicted speed respectively, both of which are three-dimensional vectors. The second-order integral and first-order integral of the gravitational acceleration at and The main calculation results.

[0124] Step 403: Construct a second nonlinear equation based on the track elements, the predicted position, and the predicted speed; the second nonlinear equation is an equation about the corrected position and the corrected speed.

[0125] Specifically, Newton iteration is applied to correct the predicted position and prediction speed , define the following about the correction position and correction speed The nonlinear equation, that is, the second nonlinear equation:

[0126] .

[0127] in, For correcting position and correction speed Track element error calculation function; is any diagonal matrix; is the semi-major axis; is the eccentricity; is the orbital inclination; is the right ascension of the ascending node; is the argument of perigee; is the argument of latitude.

[0128] Step 404: Use Newton iteration to solve the second nonlinear equation to obtain the corrected position and corrected velocity.

[0129] Step 405: Based on the corrected position and the corrected speed, the position to be increased, the speed to be increased, the first position control direction and the second position control direction are updated to obtain an updated position to be increased, an updated speed to be increased, an updated first position control direction and an updated second position control direction.

[0130] Specifically, first, based on Newton iteration, solve the correction position and correction speed , nonlinear equation relative The gradient matrix (also called the Jacobian matrix) of The value at for:

[0131] .

[0132] in, It is the orbit element error calculation function.

[0133] Secondly, solve the following about The linear equations of the displacement update are obtained and speed update amount :

[0134] .

[0135] in, For the predicted position and prediction speed Track element error calculation function.

[0136] Finally, update the position to be increased, the speed to be increased, the first position control direction and the second position control direction. The update formula is:

[0137] .

[0138] .

[0139] .

[0140] .

[0141] in, is the updated position to be added; The speed to be increased after the update; Control direction for the updated first position; Control direction for the updated second position; For the The first-order integral of the gravitational acceleration at time t; For the The second-order integral of the gravitational acceleration at time .

[0142] Step 5: Based on the updated engine shutdown time and the convergence judgment inequality, determine whether the iteration has converged to obtain a first judgment result.

[0143] As an optional implementation, the convergence judgment inequality is:

[0144] .

[0145] in, The engine shutdown time is obtained by using the iterative variable calculation formula; The stored engine shutdown time.

[0146] Specifically, The updated engine shutdown time determined at the initial time.

[0147] Step 6: If the first judgment result is no, update the current time to the next time and return to step 2.

[0148] Step 7: If the first judgment result is yes, the updated engine shutdown time and updated thrust direction are output to the rocket control system.

[0149] The updated thrust direction is determined according to the updated main amount of the thrust direction and the updated change amount of the thrust direction.

[0150] Step 8: The control system uses the updated engine shutdown time and the updated thrust direction to control the rocket operation, updates the current moment to the next moment, and returns to step 2. It then determines whether the rocket has reached the target orbit based on the flight status and orbital elements to obtain a second judgment result.

[0151] As an optional implementation, the thrust direction calculation formula is:

[0152] .

[0153] in, For the The thrust direction at the moment; is the main quantity in the thrust direction; For the time; is the change in thrust direction.

[0154] Step 9: If the second judgment result is yes, then end the guidance.

[0155] Step 10: If the second judgment result is no, update the current time to the next time and return to step 2.

[0156] Example 2

[0157] A computer device includes: a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement the dynamic explicit guidance method based on Newton iteration in embodiment 1.

[0158] Example 3

[0159] A computer-readable storage medium stores a computer program, which, when executed by a processor, implements the dynamic explicit guidance method based on Newton iteration in embodiment 1.

[0160] Example 4

[0161] A computer program product includes a computer program, which implements the dynamic explicit guidance method based on Newton iteration in embodiment 1 when executed by a processor.

[0162] Example 5

[0163] A computer device, which may be a database, may have an internal structure as shown in FIG. Figure 2As shown. The computer device includes a processor, a memory, an input / output interface (Input / Output, abbreviated as I / O) and a communication interface. The processor, memory and input / output interface are connected via a system bus, and the communication interface is connected to the system bus via the input / output interface. The processor of the computer device is used to provide computing and control capabilities. The memory of the computer device includes a non-volatile storage medium and an internal memory. The non-volatile storage medium stores an operating system, a computer program and a database. The internal memory provides an environment for the operation of the operating system and computer program in the non-volatile storage medium. The database of the computer device is used to store pending transactions. The input / output interface of the computer device is used to exchange information between the processor and an external device. The communication interface of the computer device is used to communicate with an external terminal via a network connection. When the computer program is executed by the processor, the dynamic explicit guidance method based on Newton iteration in Example 1 is implemented.

[0164] It should be noted that the object information (including but not limited to object device information, object personal information, etc.) and data (including but not limited to data used for analysis, stored data, displayed data, etc.) involved in the present invention are all information and data authorized by the object or fully authorized by all parties, and the collection, use and processing of relevant data must comply with the relevant laws, regulations and standards of relevant countries and regions.

[0165] Those skilled in the art will appreciate that all or part of the processes in the above-described method embodiments can be implemented by instructing the relevant hardware through a computer program. The computer program can be stored in a non-volatile computer-readable storage medium. When executed, the computer program can include the processes of the above-described method embodiments. Any reference to memory, database, or other media used in the embodiments provided herein may include at least one of non-volatile and volatile memory. Non-volatile memory may include read-only memory (ROM), magnetic tape, floppy disk, flash memory, optical memory, high-density embedded non-volatile memory, resistive random access memory (ReRAM), magnetic random access memory (MRAM), ferroelectric random access memory (FRAM), phase change memory (PCM), graphene memory, etc. Volatile memory may include random access memory (RAM) or external cache memory, etc. By way of illustration and not limitation, RAM can take various forms, such as static random access memory (SRAM) or dynamic random access memory (DRAM). The databases involved in the various embodiments provided herein may include at least one of a relational database and a non-relational database. Non-relational databases may include, but are not limited to, blockchain-based distributed databases. The processors involved in the various embodiments provided herein may be, but are not limited to, general-purpose processors, central processing units (CPUs), graphics processing units (GPUs), digital signal processors (DSPs), programmable logic devices (PLDs), quantum computing-based data processing logic devices, and the like.

[0166] The technical features of the above embodiments can be combined arbitrarily. To make the description concise, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0167] This document uses specific examples to illustrate the principles and implementation methods of the present invention. The above examples are only intended to help understand the method and core concept of the present invention. At the same time, those skilled in the art will find that the specific implementation methods and application scopes may vary based on the concept of the present invention. In summary, the contents of this specification should not be construed as limiting the present invention.

Claims

1. A dynamic explicit guidance method based on Newton iteration, characterized in that: The method comprises: Obtaining initial values ​​of orbital elements and iterative variables of the target orbit of the rocket; the orbital elements include: semi-major axis, eccentricity, orbital inclination, right ascension of ascending node, and argument of perigee; and the iterative variables include: position to be increased, velocity to be increased, first position control direction, second position control direction, main quantity of thrust direction, change in thrust direction, and engine shutdown time; Obtaining the flight status of the rocket; the flight status includes: position, speed, mass, thrust magnitude and effective exhaust velocity; At the current moment, using Newton iteration, based on the flight status, the main quantity of the thrust direction, the change in the thrust direction, and the engine shutdown time are updated to obtain an updated main quantity of the thrust direction, an updated change in the thrust direction, and an updated engine shutdown time; Using numerical integration and Newton iteration, based on the flight state, the updated main quantity of the thrust direction, the updated change in the thrust direction, the updated engine shutdown time, and the orbital elements, the position to be increased, the speed to be increased, the first position control direction, and the second position control direction are updated to obtain an updated position to be increased, an updated speed to be increased, an updated first position control direction, and an updated second position control direction; Based on the updated engine shutdown time and convergence judgment inequality, judging whether the iteration has converged to obtain a first judgment result; If the first judgment result is no, the current moment is updated to the next moment, and "get the rocket's flight status" is returned; If the first judgment result is yes, outputting the updated engine shutdown time and the updated thrust direction to the control system of the rocket; the updated thrust direction is determined according to the updated main quantity of the thrust direction and the updated change quantity of the thrust direction; The control system controls the operation of the rocket using the updated engine shutdown time and the updated thrust direction, updates the current time to the next time, returns to "obtaining the rocket's flight status," and determines whether the rocket has reached the target orbit based on the flight status and orbital elements, obtaining a second determination result. If the second judgment result is yes, then end the guidance; If the second judgment result is no, the current moment is updated to the next moment, and the "get the rocket's flight status" is returned.

2. The dynamic explicit guidance method based on Newton iteration according to claim 1, characterized in that: Get the initial values ​​of the rocket's iteration variables, including: Using the least propellant as the performance indicator, trajectory optimization software is used to optimize the rocket trajectory before launch to determine the position curve, velocity curve, mass curve and thrust curve, position co-state curve, velocity co-state curve and engine shutdown time; Based on the position curve, velocity curve, mass curve and thrust curve, position co-state curve, velocity co-state curve and engine shutdown time, the initial value of the rocket's iterative variable is obtained using the iterative variable calculation formula.

3. The dynamic explicit guidance method based on Newton iteration according to claim 2, characterized in that: Iteration variable calculation formula, including: ； ； ； ； ； ； ； in, For the position to be added; The engine shutdown time obtained by optimizing the rocket trajectory; For the time; For the The thrust of the moment; For the the quality of the moment; For the Velocity covariance at each moment; is the speed to be increased; Control direction for the first position; The position of the engine shutdown time obtained for optimizing the rocket trajectory; The speed of the engine shutdown time obtained for optimizing the rocket trajectory; Control direction for the second position; is the main quantity in the thrust direction; is the velocity co-state at the initial moment; is the change in thrust direction; The engine shutdown time is obtained by using the iterative variable calculation formula.

4. The dynamic explicit guidance method based on Newton iteration according to claim 1, characterized in that: Using Newton iteration, the main quantity of thrust direction, the change in thrust direction, and the engine shutdown time are updated based on the flight status to obtain the updated main quantity of thrust direction, the updated change in thrust direction, and the updated engine shutdown time, including: Constructing a first nonlinear equation based on the flight state; the first nonlinear equation is an equation about a main quantity of the thrust direction, a change in the thrust direction, and an engine shutdown time; Solving the first nonlinear equation using Newton iteration to obtain an updated amount of a main quantity in the thrust direction, an updated amount of a change in the thrust direction, and an updated amount of an engine shutdown time; updating the main quantity in the thrust direction based on the main quantity update amount in the thrust direction to obtain an updated main quantity in the thrust direction; updating the change amount of the thrust direction based on the change amount update amount of the thrust direction to obtain an updated change amount of the thrust direction; The engine shutdown time is updated based on the engine shutdown time update amount to obtain an updated engine shutdown time.

5. The dynamic explicit guidance method based on Newton iteration according to claim 1, characterized in that: By using numerical integration and Newton iteration, based on the flight state, the updated main quantity of the thrust direction, the updated change in the thrust direction, the updated engine shutdown time, and the orbital elements, the position to be increased, the speed to be increased, the first position control direction, and the second position control direction are updated to obtain the updated position to be increased, the updated speed to be increased, the updated first position control direction, and the updated second position control direction, including: Constructing a differential equation based on the flight state, the principal quantity of the updated thrust direction, and the change quantity of the updated thrust direction; Applying the fourth-order Runge-Kutta method to solve the differential equation to obtain the predicted position and predicted velocity; Constructing a second nonlinear equation based on the track element, the predicted position, and the predicted speed; wherein the second nonlinear equation is an equation about the corrected position and the corrected speed; Using Newton iteration to solve the second nonlinear equation to obtain the corrected position and the corrected velocity; Based on the corrected position and the corrected speed, the position to be increased, the speed to be increased, the first position control direction and the second position control direction are updated to obtain an updated position to be increased, an updated speed to be increased, an updated first position control direction and an updated second position control direction.

6. The dynamic explicit guidance method based on Newton iteration according to claim 1, characterized in that: The convergence judgment inequality is: ； in, The engine shutdown time is obtained by using the iterative variable calculation formula; The stored engine shutdown time.

7. The dynamic explicit guidance method based on Newton iteration according to claim 1, characterized in that: The formula for calculating the thrust direction is: ； in, For the The thrust direction at the moment; is the main quantity in the thrust direction; For the time; is the change in thrust direction.

8. A computer device comprising: A memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement the dynamic explicit guidance method based on Newton iteration according to any one of claims 1 to 7.

9. A computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the dynamic explicit guidance method based on Newton iteration described in any one of claims 1 to 7 is implemented.

10. A computer program product comprising a computer program, characterized in that When the computer program is executed by a processor, the dynamic explicit guidance method based on Newton iteration described in any one of claims 1 to 7 is implemented.

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